VTOL aircraft

ABSTRACT

The aircraft can include: an airframe, a tilt mechanism, a payload housing, and can optionally include an impact attenuator, a set of ground support members (e.g., struts), a set of power sources, and a set of control elements. The airframe can include: a set of rotors and a set of support members.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of U.S. Provisional Application No.62/838,773, filed 25 Apr. 2019, which is incorporated in its entirety bythis reference. This application claims the benefit of US ProvisionalApplication No. 62/983,445, filed 28 Feb. 2020, which is incorporated inits entirety by this reference.

This application is related to U.S. application Ser. No. 16/708,280,filed 9 Dec. 2019, U.S. application Ser. No. 16/430,163, filed 3 Jun.2019, and U.S. application Ser. No. 16/409,653, filed 10 May 2019, eachof which is incorporated in its entirety by this reference.

TECHNICAL FIELD

This invention relates generally to the aviation field, and morespecifically to a new and useful aircraft in the aviation field.

BRIEF DESCRIPTION OF THE FIGURES

FIG. 1A is a schematic representation a variant of the system in thehover arrangement from a top view.

FIG. 1B is a schematic representation the variant of the system in FIG.1A in the forward arrangement from a top view.

FIG. 1C is a side view of the variant of the system in FIG. 1A in thehover arrangement.

FIG. 1D is a side view of the variant of the system in FIG. 1B in theforward arrangement.

FIG. 2 is a schematic representation of a variant of the systemtransforming a rotor between a forward configuration and a hoverconfiguration.

FIG. 3A and FIG. 3B are schematic representations of a variant of thesystem from a top view and a side view, respectively, illustratingvarious axes of the aircraft.

FIGS. 4A-4H are each a schematic representation a different variant ofthe airframe in the forward arrangement from a front view.

FIGS. 5A-5F are each a schematic representation a different variant ofthe rotor arrangement in the hover arrangement from a top view.

FIG. 6A is a side view schematic representation of a variant of theaircraft in the forward arrangement with the lift vector axis aligned tothe weight vector and with the forward thrust vector axis aligned to thedrag vector.

FIG. 6B is a side view schematic representation of a variant of theaircraft in the hover arrangement with the lift vector axis aligned tothe weight vector.

FIG. 7 is a side view representation of a variant of a rotor.

FIGS. 8A, 8B, and 8C are side view representations of a variant with ananti-lateral support member respectively forward, behind, andintersecting a lateral support member.

FIG. 9 is a flow chart diagram of a variant of the method.

FIGS. 10A and 10B are a front view schematic representations of a firstand second variant of the system in the forward arrangement,respectively.

FIG. 11A is a side view schematic representation of a variant of apayload housing including a passenger region.

FIG. 11B is a side view schematic representation of a variant of apayload housing including a passenger region.

FIG. 12A is a top view schematic representation of a variant of theaircraft including a rear rotor in the hover configuration.

FIG. 12B is a top view schematic representation of a variant of theaircraft including a rear rotor in the forward configuration.

FIG. 12C is a side view schematic representation of a variant of theaircraft including a rear rotor in the hover configuration.

FIG. 12D is a side view schematic representation of a variant of theaircraft including a rear rotor in the forward configuration.

FIGS. 13A-C are schematic representations of a variant of the rotor discangle of attack relative to a wing in a forward configuration,transition configuration, and hover configuration, respectively.

FIGS. 14A, 14B, and 14C are a top view schematic representation of avariant of the aircraft in a hover configuration, a top view schematicrepresentation of the variant in a forward configuration, and a sideview schematic representation of the variant in a hover configuration,respectively.

FIGS. 15A, 15B, and 15C are a top view schematic representation of avariant of the aircraft in a hover configuration, a top view schematicrepresentation of the variant in a forward configuration, and a sideview schematic representation of the variant in a hover configuration,respectively.

FIGS. 16A, 16B, and 16C are a top view schematic representation of avariant of the aircraft in a hover configuration, a top view schematicrepresentation of the variant in a forward configuration, and a sideview schematic representation of the variant in a hover configuration,respectively.

FIGS. 17A, 17B, and 17C are a top view schematic representation of avariant of the aircraft in a hover configuration, a top view schematicrepresentation of the variant in a forward configuration, and a sideview schematic representation of the variant in a hover configuration,respectively.

FIG. 18A-F are front view schematic representations of a variant of theaircraft which include dihedral wings.

FIG. 19A-F are front view schematic representations of a variant of theaircraft which include anhedral wings.

FIGS. 20A-D are schematic representations of example rotor rotationdirections.

FIG. 21 is a side view schematic representation of an example payloadhousing including an insulated cargo region.

FIG. 22A is a side view cross sectional representation of a variant ofthe airframe, illustrating a connection between lateral and anti-lateralsupport members.

FIG. 22B is a partial schematic representation of a variant of theairframe.

FIG. 22C is a schematic representation of a variant of the airframe.

FIG. 23A-B are isometric views of a variant of the aircraft in theforward and hover configurations, respectively.

FIG. 24A-B are isometric views of a variant of the aircraft in theforward and hover configurations, respectively.

FIG. 25A-B are top views of a variant of the aircraft in the hover andforward configurations, respectively.

FIG. 25C-D are front views of a variant of the aircraft in the hover andforward configurations, respectively.

FIG. 25E-F are side views of a variant of the aircraft in the hover andforward configurations, respectively.

FIG. 25G is a side view of a variant of the aircraft in the forwardconfiguration.

DESCRIPTION OF THE PREFERRED EMBODIMENTS

The following description of the preferred embodiments of the inventionis not intended to limit the invention to these preferred embodiments,but rather to enable any person skilled in the art to make and use thisinvention.

1. Overview

The aircraft 100 can include: an airframe 101, a tilt mechanism 110, apayload housing 120, and can optionally include an impact attenuator150, a set of ground support members (e.g., struts) 160, a set of powersources 170, and a set of control elements 180. The airframe caninclude: a set of rotors 130 and a set of support members 140. However,the aircraft 100 can additionally or alternatively include any othersuitable set of components. A first example of the aircraft 100 is shownin FIGS. 1A-1D. A second example of the aircraft is shown in FIGS.25A-F.

The term “rotor” as utilized herein, in relation to the aircraft orotherwise, can refer to a rotor, a propeller, and/or any other suitablerotary aerodynamic actuator. While a rotor can refer to a rotaryaerodynamic actuator that makes use of an articulated or semi-rigid hub(e.g., wherein the connection of the blades to the hub can bearticulated, flexible, rigid, and/or otherwise connected), and apropeller can refer to a rotary aerodynamic actuator that makes use of arigid hub (e.g., wherein the connection of the blades to the hub can bearticulated, flexible, rigid, and/or otherwise connected), no suchdistinction is explicit or implied when used herein, and the usage of“rotor” can refer to either configuration, and any other suitableconfiguration of articulated or rigid blades, and/or any other suitableconfiguration of blade connections to a central member or hub. Likewise,the usage of “propeller” can refer to either configuration, and anyother suitable configuration of articulated or rigid blades, and/or anyother suitable configuration of blade connections to a central member orhub. Accordingly, a tiltrotor aircraft can be referred to as atilt-propeller aircraft, a tilt-prop aircraft, and/or otherwise suitablyreferred to or described.

The term “center of gravity” (CoG) as utilized herein, in relation tothe aircraft 100 or otherwise, can refer to a point from which theweight of the body or system may be considered to act, and isinterchangeable with the term “center of mass” (CoM) herein (e.g., underassumption of substantially uniform gravity). The CoG of the aircraftcan refer to the aircraft CoG in any suitable state and/orconfiguration: loaded and/or unloaded; forward, transition, and/orvertical configuration; airframe without and/or without payload housingattached; and/or any other suitable aircraft state or configuration.

The term “center of lift” (CoL) as utilized herein can refer to thepoint where the sum total of all lift generated by aircraftparts—principally by wings, rotors, control surfaces, and/or aerodynamicfuselage parts (e.g., exterior of cargo housing) but additionally oralternatively by other aircraft components—generates a net moment ofzero about the CoL and the aggregate lift force (e.g., collectivelygenerated by the aircraft components) will act through the CoL while inan atmosphere. The CoL can be the same as the CoG, collocated with theCoG, forward of the CoG, rearward of the CoG, outboard of the CoG,inboard of the CoG, and/or otherwise positioned relative to the CoG. TheCoL (and/or the location of the CoL relative to the CoG) can becontrollable and/or adjustable by control of the rotors, tilt mechanism,control surfaces, shifting the CoG to account for differentdistributions of mass (e.g., passengers, cargo, fuel, etc.) onboard theaircraft, and/or otherwise controlled. In a first specific example, thecontrol of the rotors is adjusted such that the CoL is substantially thesame as the CoG (or otherwise located on the vertical axis), resultingin stable forward flight.

The term “aerodynamic center,” in reference to a rotor, wing, airfoil,or otherwise, can refer to the point around which the aerodynamicmoments do not change with changes in aircraft attitude. The aerodynamiccenter can be in the same location as the CoL or can be in a differentlocation from the CoL. References to the CoL hereinafter can be equallyapplicable to the aerodynamic center, or be treated differently.

The term “geometric center of the rotors” can refer to an absolute orrelative point (e.g., relative to the airframe across all positions ofthe tilt axis) which minimizes the sum of distances to the centers ofall the rotors (e.g., rotor hubs) in 3-space or in a projected plane(e.g., vertical/lateral plane, frontal plane, top plane, etc.).Alternately, the geometric center of the rotors can refer to the averagelocation of the rotors, a point equidistant to each rotor pair, and/orcan be otherwise suitably defined.

The term “center of thrust” (CoT) as utilized herein can refer to thelocation at which the resultant or total thrust can be taken to act(magnitude and direction, the latter being sometimes referred to as the‘thrust line’). The CoT can be controllable and/or adjustable by controlof the rotors, tilt mechanism, and/or drag-inducing components (e.g.,cargo housing, landing gear, etc.). In a first specific example, the CoTacts through the CoG of the aircraft (e.g., is aligned with the CoG),and there is no resulting moment causing the aircraft to pitch, yaw, orroll. In a second specific example, the CoT does not act through the CoGof the aircraft, and there is a resulting moment which will cause theaircraft to pitch, yaw, or roll (unless counteracted by another aircraftmoment).

The term “substantially” as utilized herein can mean: exactly,approximately, within a predetermined threshold or tolerance, and/orhave any other suitable meaning.

In examples, components of the systems and method described herein canbe used, replaced, and/or combined with the aircrafts described in U.S.application Ser. No. 14/218,845, filed 14 Mar. 2014 and/or U.S.application Ser. No. 14/662,085, filed 18 Mar. 2018. However, thesystems and methods can be otherwise configured.

2. Benefits

Variations of the technology can afford several benefits and/oradvantages.

First, variations of this technology can generate lift with the rotorsof the aircraft in a forward and/or transition flight mode. By utilizinga larger rotor blade area (and/or larger rotor disc area) and adjustingthe blade pitch and RPM, the rotors can augment the lift generated bythe aerodynamic profile of the aircraft in the forward flight mode inaddition to providing forward thrust. Generating lift with the rotorscan augment the lift generated by the aerodynamic profile of theaircraft (e.g., wings/support member geometry, fuselage geometry, etc.)or otherwise generate sufficient lift to support the aircraft in aforward configuration.

Second, variations can utilize shorter and/or stiffer support members(e.g., wings), which can improve the handling characteristics of theaircraft (e.g., tighter turning radius, smaller required landing zone,etc.). Variants with shorter/stiffer support members reduce the risk ofwhirl flutter, which can weaken and/or destroy the wings, and results inan unsatisfactory/dangerous ride experience. Reducing the risk of whirlflutter can be particularly beneficial in variants utilizing a smallerwing area relative to the total rotor area, as smaller airfoil crosssections can result in a reduction in rigidity. Variants withshorter/stiffer support members can reduce the total profile of theaircraft (e.g., size of the aircraft—particularly the width), can belower cost, and can reduce the drag of the aircraft.

Third, variants generating lift with the rotors can reduce or eliminateadditional control surfaces (e.g., wing flaps, ailerons, ruddervators,elevators, rudder, etc.) on the aircraft since the thrust and motortorque is controllable (thereby indirectly controlling lift) at eachrotor, thereby enabling pitch, yaw, and/or roll control during forwardflight. In some variants, the aircraft is not required to bank to turnin the forward configuration because a set of anti-lateral supportmembers and/or rotors can generate lateral aerodynamic forces (e.g., insideslip) and/or net yaw moments to change the heading of the aircraft.The first, second, and third specific variants can independently orcollectively reduce and/or eliminate any suitable set of controlsurfaces on the aircraft, enable lift generation with rotors, headingchanges without banking the aircraft, and/or elimination of any suitableset of control surfaces. In a first specific example, full controlauthority can be achieved about axes perpendicular to the average rotoraxis—angular accelerations around those axes can be achieved throughthrust redistribution of rotors. Along the predominant rotor axis,rotation can be achieved via redistribution of torque to individualmotors such that the desired total axial thrust is preserved.Additionally, if rotor axes are slightly canted relative to each other(e.g., different angles of attack, different angles of attack relativeto the wing), redistribution of thrust amongst them can also createadditional moments along the average axis of all the rotors. However,the rotors can otherwise suitably generate lift.

Fourth, variations of this technology can utilize a large total rotordisc area relative to the width of the aircraft and/or a large rotorblade area relative to the wing area, which can reduce the acousticprofile of the aircraft. Utilizing a large blade area for eachindividual rotor (e.g., for a given rotor disc diameter) allows for lowrotor tip speeds (e.g., relative to tip speeds of a rotor of theeffective rotor disc diameter, Mach 0.3, etc.), which can reduce theacoustic profile for the aircraft. Additionally, large rotor bladesareas for each individual rotor can enable rotors to operate moreefficiently for lift generation during forward flight (as discussedabove). However, the aircraft can otherwise suitably reduce the acousticprofile and/or include other suitable rotors.

Fifth, variants of the technology utilize a non-structural payloadhousing, which allows the payload housing to be lighter and lower cost.In such variants, support members (and/or torsion boxes) can be arrangedabove and/or behind the payload housing to avoid infringing on thepayload housing space providing more space for passengers/cargo andreducing/eliminating the need for additional structural support in thepayload housing. In variants utilizing a non-structural payload housing,the component count in the payload housing can be reduced, as batteries,the primary electrical architecture, control systems, sensors, and othercomponents can be moved to other parts of the aircraft (e.g., wings,support members, nacelles, etc.). In variants, the payload housing canbe non-structural because it does not need to support landing gearloads, because the aircraft does not utilize conventional landing gear.Instead, struts on the nacelles (which extend to the ground in the hoverconfiguration) provide a reliable, light weight, and low cost means tosupport the aircraft on the ground, and can further provide aerodynamicadvantages in forward flight because of their streamlined geometry.Struts on the nacelles directly transfer landing gear loads to thesupport members (e.g., airframe) without directing them through thepayload housing. In variants, the payload housing can be non-structuralbecause the support structure of the airframe (e.g., the set ofstructural members) can provide the structural rigidity and attachmentpoints for the aircraft components. However, non-structural payloadhousings can be otherwise achieved. In variants of the technologyutilizing a non-structural payload housing, the payload housing can bemodular, detachable, and/or reconfigurable to allow fasterloading/unloading of cargo (e.g., passengers, cargo, etc.), which canimprove the uptime of the aircraft. Additionally, modular, detachable,and/or reconfigurable payload housings can enable additional means ofground transport of a pod, such as towing, vehicular transport, or othermodes of module transportation. Further, variants utilizing a modularand/or reconfigurable payload housing option can enable switchingaircraft frames for faster charging and/or refueling to further improveoperational efficiency.

Sixth, variants of the technology can minimize points of failure becausethe aircraft is capable of landing in any orientation of the tiltmechanism, with one or more rotors inoperable, and/or with one or morecontrol surfaces inoperable. In a specific example, if the tiltmechanism is stuck and/or locked between the forward and hoverconfigurations, the aircraft can still land by reorienting the rotorsvertically (e.g., by controlling the pitch of the aircraft upwards) withthe payload housing at an angle (e.g., skewed/pitched upwards) relativeto the ground. In variants, this is possible due to the number anddistribution of redundant rotors on the aircraft, but can be otherwiseachieved. Variants of the technology can utilize an impact attenuator toensure that passengers and/or cargo are protected when landing in askewed orientation of the payload housing (e.g., if the tilt mechanismfails between the forward configuration and the hover configuration).

Seventh, variants of the technology offer improved stability, trim,and/or maneuverability in all modes of flight. Variants can achieve thisstability by axis alignment of the center of (forward) thrust with thecenter of drag (and/or gravity) and the center of lift with the centerof gravity in the forward flight mode, and axis alignment of the centerof lift (e.g., vertical rotor thrust) with the weight vector in thehover mode. Axis alignment of one or more axes can be achieved bytrimming (e.g., automatically) the thrust and/or lift distribution ofthe rotors, such as by power distribution between rotors, actuation ofrotor blades, and/or other suitable control. In variants, the geometriccenter of the rotors (e.g., average hub location) is substantiallyaligned with the (forward) thrust and/or center of drag in the forwardflight mode (e.g., with or without selective power provision), so as tominimize the trimming required, enable even power distribution, and/ormaintain control authority (e.g., required power redistribution does notexceed a threshold, power redistribution is within continuous operationregime of motors). In other variants, the geometric center of the rotorscan be misaligned with the center of mass; in these variants, the rotorscan be selectively powered (e.g., power selectively redistributed) toadjust the thrust and/or lift distribution of the rotors into thetrimmed condition. In specific examples, the rotor placement can beselected to substantially equally distribute power across the rotorsdespite this misalignment (e.g., to maintain control authority andefficiency; minimize trimming), or be otherwise arranged. Additionally,variants including smaller/stiffer support members (e.g., wings) offerimproved handling characteristics because they can perform tighter turnsand require a smaller area in order to land. In some variants, theaircraft is not required to bank in order to turn during forward (andhover/transition) flight modes, which can improve handling and/or ridecomfort for passengers.

However, variations of the technology can additionally or alternatelyprovide any other suitable benefits and/or advantages.

3. System

The aircraft 100 can include: an airframe, a tilt mechanism, a payloadhousing, and can optionally include an impact attenuator, a set ofground support members (e.g., struts), a set of power sources, and a setof control elements. However, the aircraft 100 can additionally includeany other suitable set of components.

The aircraft 100 can be to any suitable type of aircraft. The aircraft100 is preferably a tilt-wing aircraft (e.g., other similarly configuredaircraft which tilt lateral support members relative to the payloadhousing), but can additionally or alternately be a tilt-rotor aircraft,rotorcraft, propeller aircraft, fixed wing aircraft, lighter-than-airaircraft, heavier-than-air aircraft, and/or any other suitable aircraft.The aircraft can operate as: VTOL, STOL, STOVL, takeoff like a fixedwing aircraft, land like a fixed wing aircraft, and/or operate in anyother suitable manner. The aircraft can be manned, unmanned (e.g.,autonomous, remotely piloted, etc.), a cargo aircraft, passengeraircraft, drone, and/or other suitable type of aircraft. The aircraft ispreferably operable between a forward configuration, hoverconfiguration, and transition configuration (e.g., between forward andhover), but can additionally or alternately be operable in a taxi (e.g.,ground operation) configuration, and/or be otherwise suitablyconfigured. An example of the forward and hover configurations is shownin FIG. 2.

The aircraft 100 defines various geometrical features. The aircraftdefines principal geometric axes, as shown in FIGS. 3A-3B, including: avertical axis 105 (e.g., yaw axis), a longitudinal axis 104 (e.g., aroll axis), and a lateral axis 103 (e.g., a pitch axis). The vertical,longitudinal, and lateral axes can be defined such that they intersectat the center of gravity (CoG) of the aircraft 102, and a pure momentabout any one of the aforementioned axes causes the aircraft 100 torotate about the vertical, longitudinal, and lateral axes, respectively.However, the three principal axes can additionally or alternatively bedefined geometrically (e.g., based on lines of symmetry of the aircraftin one or more dimensions, based on arbitrary lines through theaircraft, etc.) with or without reference to the CoG. For example, theaxes can intersect at a geometric center of the aircraft. The propellersof the aircraft each define a disc area centered at the axis of rotationof the propeller, and the disc area is contained by an infinite discplane extending away from the axis of rotation. In variations of theaircraft, the disc planes of each of the plurality of rotors can becoextensive with any suitable subset of the remainder of the pluralityof propulsion assemblies. In a first example, each disc plane can becoextensive with each other disc plane in the hover configuration of afirst variation. In a second example, each disc plane can be coextensivewith the disc plane of one other propulsion assembly symmetricallyacross the longitudinal axis of the aircraft and displaced from (e.g.,offset from) the disc planes of each other propulsion assembly.Propeller axes can be coaxial with motor axes and/or other propelleraxes, not coaxial with motor axes and/or other propeller axes, coplanar,not coplanar, and/or otherwise suitably oriented relative to motor axesand/or other propeller axes. However, the propeller axes and/or discplanes of the plurality of propulsion assemblies can be otherwisesuitably arranged relative to one another.

The aircraft 100 can operate within any suitable acoustic range. Theaircraft preferably operates below a maximum dB level, but canadditionally or alternately operate within different acoustic ranges indifferent flight configurations (e.g., forward, transition, hover,taxi), and/or be configured to operate in different acoustic ranges(e.g., such as when near human populations, urban centers, to complywith varying regulatory restrictions, etc.). Variants can utilize alarge total rotor disc area relative to the width of the aircraft and/ora large rotor blade area relative to the wing area, which can reduce theacoustic profile of the aircraft. In variants, the individual rotor discdiameter is between 10% and 40% of the width of the aircraft, but can be10%, 20%, 30%, 40%, >40%, and/or <10% of the width of theaircraft—employing a plurality of such rotors can enable an effectiverotor disc diameter (e.g., single theoretical disc of the same area asthe total combined rotor disc area of the individual rotors) of >40% thewidth of the aircraft, such as <40%, 50%, 60%, 70%, 80%, 90%, 100%,120%, 150%, 200%, 250%, and/or >250%. Similarly, utilizing a large bladearea (e.g., overall exposed surface area, blade platform area, etc.) foreach individual rotor (e.g., for a given rotor disc diameter) allows forlow rotor tip speeds (e.g., relative to tip speeds of a rotor of theeffective rotor disc diameter, Mach 0.3), which can reduce the acousticprofile for the aircraft. The individual rotor blade area (e.g., definedas the chord integrated in the radial direction of the rotor disc andmultiplied by the number of blades on the rotor, the integrated chordalong the radius of a blade, or otherwise defined) and/or total rotorblade area (e.g., the sum of all the individual rotor blade areas for arotor, for the vehicle, etc.) is preferably between 10% and 200% of thewing area, but can be <5%, 10%, 20%, 30% 50%, 75%, 100%, 150%,200%, >200%, ranges therebetween, and/or any other suitable proportionrelative to the wing area. In a specific example, the total rotor bladearea can be 10 square meters. In a second specific example, the wingarea can be 8 square meters and the wing span can be 10 meters. In suchvariants, the aircraft can be configured to operate within an acousticrange in the hover mode with a minimum dB level of: less than 30, 40,50, 55, 60, 65, 70, 75, 80, 85, 90, 95, 100, 105, 110, or any othersuitable dB level; and a maximum dB level of 40, 45, 50, 55, 60, 65, 70,75, 80, 85, 90, 95, 100, 105, 110, 115, 120, more than 120, or any othersuitable dB level. In such variants, the aircraft can be configured tooperate within an appropriate acoustic range in the forward mode with aminimum dB level of: less than 10, 30, 40, 50, 55, 60, 65, 70, 75, 80,85, 90, 95, 100, 105, 110, or any other suitable dB level; and a maximumdB level of 40, 45, 50, 55, 60, 65, 70, 75, 80, 85, 90, 95, 100, 105,110, 115, 120, more than 120, ranges therebetween, and/or any othersuitable dB level. The acoustic range can similarly be determined bytransforming this acoustic range into an EPNL scale (EPNdB), A-weighted(dBA), C-weighted (dBC), Z-weighted, CNEL, NDL, SEL, SENEL, Leq, Lmax,and/or other expression of noise level, measured at a distance of 0 m,10 m, 25 m, 50 m, 100 m, 150 m, 200 m, 300 m, 500 m, 1000 m, and/or anyother appropriate proximity; alternatively, the numbers discussed abovefor the acoustic range can be applied to the aforementioned noise levelexpressions.

The aircraft 100 can have any suitable mass and/or mass limitations(e.g., unloaded mass, loaded mass, max takeoff mass, etc.). The aircraftmass can be: 0.1 kg, 0.25 kg, 0.5 kg, 0.75 kg, 1 kg, 2 kg, 3 kg, 5 kg,10 kg, 50 kg, 200 kg, 500 kg, 1000 kg, 1250 kg, 1500 kg, 1750 kg, 2000kg, 2250 kg, 2500 kg, 3000 kg, 3500 kg, 5000 kg, 10000 kg, 25000 kg,greater than 25000 kg, less than 0.1 kg, less than 1 kg, 1-5 kg, 5-10kg, less than 10 kg, less than 1000 kg, less than 3500 kg, less than10000 kg, greater than 25000 kg, and/or any other suitable mass. Thecargo/payload mass capacity can be: 0.1 kg, 0.25 kg, 0.5 kg, 0.75 kg, 1kg, 2 kg, 3 kg, 5 kg, 10 kg, 50 kg, 200 kg, 500 kg, 1000 kg, and/or anyother suitable mass. The fuel and/or battery mass can be: 0.1 kg, 0.25kg, 0.5 kg, 0.75 kg, 1 kg, 2 kg, 3 kg, 5 kg, 10 kg, 50 kg, 200 kg, 500kg, 1000 kg, and/or any other suitable mass. The fuel and/or batterycapacity can define any suitable aircraft range, which can be <1 mi, 1mi, 5 mi, 10 mi, 20 mi, 50 mi, 100 mi, 150 mi, 200 mi, 250 mi, and/orany other suitable aircraft range.

In variants, the rotors generate lift by angling (e.g., actively,passively such as by an installation angle, etc.) the thrust axis one ormore rotors of the aircraft by 5-7 degrees, generating a lift forceperpendicular to the direction of incoming airstream, with a certainincrease in necessary power to maintain the same thrust. The ratio oflift force multiplied by forward velocity to the power needed tomaintain 0 forward thrust is the L/De (lift over equivalent drag)efficiency merit—this can be on the order of 20 for structurally supportand/or transform aerodynamic aircraft components (e.g., rotors, wings,control surfaces, etc.). The airframe can optionally provide attachmentpoints to modular component attachment, such as payload housingattachment. The aircraft can include one or more airframes. Eachairframe preferably includes a support structure, but can additionallyor alternatively include any other suitable component. The supportstructure can include a set of support members, and can function tomount the set of rotors (e.g., between the forward and hover modes), themodular components, the tilt mechanism(s), and/or any other suitablecomponent.

The rotors of the aircraft 100, in variants, function to provide thrust(e.g., force aligned with direction of motion, force propelling aircraftin direction of motion) and/or lift (e.g., force opposing gravity, forceorthogonal to thrust, etc.). The rotors preferably provide lift tosustain flight of the aircraft which is equal to or greater than theaircraft weight in all flight configurations, however the rotors canadditionally or alternately provide a portion of the required lift inthe flight configurations, including forward (and/or transition)configurations (e.g., in conjunction with a set of wings). An examplerotor is shown in FIG. 7. The rotors can provide, individually orcollectively: 300%, 200%, 150%, 120%, 110%, 95%, 90%, 75%, 50%, 40%,30%, 20%, 10%, 5%, more or less than the aforementioned percentages, orany suitable proportion of the lift required to maintain aircraftaltitude during forward flight or the total lift generated by theaircraft during flight. The rotors can additionally or alternativelyprovide, individually or collectively: 500%, 250%, 150%, 120%, 110%,95%, 90%, 75%, 50%, 40%, 30%, 20%, 10%, 5%, or any suitable proportionof the force required to propel the aircraft during forward flight,liftoff (e.g., takeoff), transitional operation, and/or any othersuitable operation mode. The rotors can additionally or alternativelyprovide, individually or collectively: 300%, 150%, 120%, 110%, 95%, 90%,75%, 50%, 40%, 30%, 20%, 10%, 5%, or any suitable proportion of the liftand/or thrust required to propel the aircraft during hover flight.

The rotors can provide lift associated with an angle of attack of therotor disc relative to: the incoming airstream during forward flight,longitudinal axis of the aircraft, wing of the aircraft (e.g., the chordline 146 of the wing cross section), and/or other reference axis orplane. The angle of attack of the rotor disc relative to the wing (e.g.,chord line 146 of the wing) can be <0 deg, 0 deg, 1 deg, 3 deg, 5 deg, 7deg, 9 deg, 11 deg, 13 deg, 15 deg, any range bounded by theaforementioned values, and/or any other suitable angle. An example ofthe angle of attack of the rotor disc relative to the wing 191 isillustrated in FIG. 13A. The angle of attack of the rotor disc can be <0deg, 0 deg, 1 deg, 3 deg, 5 deg, 7 deg, 9 deg, 11 deg, 13 deg, 15 deg,17 deg, 20 deg, >20 deg, any range bounded by the aforementioned values,and/or any other suitable angle. An example of the angle of attack ofthe rotor disc 192 is illustrated in FIG. 13B. The rotor disc angle ofattack (relative to the wing or otherwise) can be defined (e.g.,measured) relative to the rotor axis of rotation, motor axis ofrotation, a vector orthogonal to the rotor disc plane 133, and/or anyother suitable reference. The angle of attack of the rotor preferablytransforms based on the transformation of the tilt mechanism and/orpitch of the aircraft. Preferably, the rotor disc planes aresubstantially parallel to the lateral/longitudinal plane (pitch/rollplane) in the hover configuration, and angled relative to thevertical/lateral plane (yaw/pitch plane) in the forward configuration(and/or hover configuration). Accordingly, the tilt mechanism preferablytransforms the wing by 90 degrees less the rotor disc angle of attackwhile transitioning between the forward and hover configurations (anexample is illustrated in FIG. 13C), however the tilt mechanism cantransform the wing by 90 degrees plus the rotor disc angle of attackwhile transitioning between the forward and hover configurations,exactly 90 degrees between the forward and hover configurations, and/orany other suitable transformation angle. In a specific variant, thetransformation between forward and hover can include tilting pastvertical (e.g., creating a rearward thrust vector) in order to arrestforward motion of the vehicle.

Each rotor preferably includes a hub, which couples the rotor blades 136to the propulsion system. The propulsion system can be rigidly coupledto the airframe (e.g., wing) structure and/or a nacelle, or can becoupled via a rotor tilt mechanism or articulated linkage configured totransform the rotor relative to the wing, change the angle of attack,and/or change a side cant angle relative to the wing. The propulsionsystem is preferably an electric motor (e.g., capable of 70 kWcontinuous power), which can be integrated in to the hub or separate anddistinct from the hub. Alternately, the propulsion system can be aninternal combustion engine (ICE), turbine engine, a hybrid-electricengine, and/or any other suitable propulsion system. In variants, one ormore rotors can be coupled and/or linked to the same propulsion systemvia shafts, rotational couplings, cross-linkages, and/or other suitablemechanisms. The hub preferably defines the axis of rotation of therotor. In a specific example, the aircraft can include a plurality ofpropulsion assemblies, each propulsion assembly including: an electricmotor and a propeller rotatably coupled to the electric motor about anaxis of rotation. In a second specific example, the hub can be locatedat the geometric center of the rotor and/or define the geometric centerof the rotor.

Each rotor can include a set of rotor blades 136, which function togenerate an aerodynamic force as they are rotated through a fluid (e.g.,air), which can be used to propel the aircraft. Each rotor can includeany suitable number of rotor blades. Preferably, each rotor includes 5rotor blades, but can alternately include 2, 3, 4, 6, or more than sixrotor blades for each rotor. The rotor blades can have any suitableblade cross section and/or aerodynamic profile. In a first specificexample, the rotor blades are the rotary airfoil blades described inU.S. application Ser. No. 16/708,280, filed 9 Dec. 2019, which isincorporated in its entirety by this reference. However, the rotorblades can be otherwise configured.

The rotor blades can define any appropriate spanwise geometry.Preferably, the upper surface of the rotor blades is generally in avesica piscis geometry, but can additionally or alternately be taperedtoward the tip (e.g., decreasing rotary airfoil chord length across theend portion of the blade), have constant cross sectional area, havevariable cross sectional area, and/or have any other appropriategeometry. The taper angle can be the same or different on the leadingedge, the trailing edge of the airfoil, on an inner portion of therotary airfoil, and/or at the tip. The tip of the rotary airfoil canhave any appropriate geometry. The tip can be flat, rounded, or pointed,and can be a point, edge, face, and/or other appropriate geometry. Therotary airfoil can have any appropriate tip angle. The blade tip can beanhedral, dihedral, un-angled, and/or at any suitable angle. The rotaryairfoil can have any appropriate twist angle. The twist angle preferablychanges the effective blade angle of attack along the span of the rotaryairfoil. The blade twist angle is preferably defined between theinnermost and outer (tip) cross sections, but can be defined between anytwo cross sections, a section of the blade, and/or at any suitableangle.

The rotor blades can have any appropriate angular spacing about the axisof rotation. Preferably, the rotor blades are evenly spaced about theaxis of rotation, but can alternately be spaced unevenly about the axisof rotation (e.g., for sound mitigation). In a first specific example,the rotor blades are spaces about the axis of rotation as described inU.S. application Ser. No. 16/430,163, filed 3 Jun. 2019, which isincorporated in its entirety by this reference. However, the rotorblades can be otherwise arranged.

The rotor blades can define a span of any appropriate length (e.g.,blade length). The span can be sized relative to a cross sectional chordlength (L), independent of the chord length, and/or any appropriatelength. The span can be: 1 L, 5 L, 10 L, 15 L, 20 L, 25 L, 50 L, <5 L,5-25 L, 25-50 L, >50 L, <5 cm, 5 cm, 10 cm, 25 cm, 30 cm, 35 cm, 40 cm,45 cm, 50 cm, 60 cm, 70 cm, 80 cm, 90 cm, 1 m, 1.25 m, 1.5 m, 1.75 m,2.5 m, 5 m, 10 m, 15 m, 20 m, 5-25 cm, 25-50 cm, 50-100 cm, 0.1 m-15 m,1-2 m, 1-4 m, 5-10 m, 10-20 m, >20 m, and/or any other suitable length.In a specific example, the rotor blades define a rotor disc diameter of3 meters.

Separate rotors on the aircraft preferably operate with the same rotorblades in the same configuration, but can alternately include adifferent: number of rotor blades, rotor blade length (or radius ofrotor disc), rotor blade spacing, rotor blade cross section, and/orother different characteristic.

The rotors can include a blade pitching mechanism 137, which functionsto change the angle of attack of the rotor blade (e.g., relative to thefluid flow). The rotor can include a single pitching mechanism, ormultiple pitching mechanisms associated with each rotor such as: one perrotor, multiple per rotor, one per blade, and/or any other suitablenumber of blade pitching mechanisms per rotor. The pitching mechanismcan actuate blades independently or actuate multiple simultaneously. Thepitching mechanism can be: integrated into the rotor hub,connected/mounted to the rotor hub, and/or separate from the rotor hub.Preferably, the pitching mechanism can be electromechanically actuated,but can additionally or alternately be hydraulic, pneumatic, and/orground adjustable (by a human operator or other input). The pitchingmechanism can be operable between a finite or infinite number ofpositions. The pitching mechanism can be: a controllable-pitch propeller(CPP), a swashplate, a ground adjustable rotor, and/or other pitchingmechanism. In a first variant, the blade pitch mechanism is aswashplate. In a second variant, the blade pitching mechanism is a setof electromechanical actuators. In a third variant, the rotor does notinclude a pitching mechanism and the lift generated by the rotors iscontrolled by varying the RPM.

The rotors can include a nacelle 138, which functions as the structuralmounting for the rotors. Nacelles can additionally function as packagingfor one or more propulsion components (e.g., motor, engine, etc.) and/orpower sources. The nacelles are preferably an aerodynamically efficientshape (e.g., teardrop), which tapers toward a trailing portion of thenacelle in the forward configuration. The nacelles preferably serve as asupport member node connecting the rotor to the airframe, and can beconnected to 1, 2, 3, or more than 3 support members. In a firstvariant, a nacelle can be connected to the endpoint of a support member.In a second variant, a nacelle can bisect a support member. In a thirdvariant, a nacelle can be directly integrated into a support memberand/or a support member can be directly integrated into the nacelle(e.g., landing gear strut, ground support member, etc.). Nacelles arepreferably fixed relative to the lateral support members (and/or wing orother mounting component), but alternately can rotate, slide, orotherwise actuate relative to the wing and/or the tilt mechanism. In aspecific example, nacelles are mounted at the outboard termination of alateral support member (e.g., left wing and/or right wing, outboard endof a wing) and the vertical termination (e.g., upper end and/or lowerend) of an anti-lateral support member. In a second specific example,the nacelles can be mounted to the leading edge or side of the supportmembers, proximal a wing extremity (e.g., wing end). However, thenacelles can have any other suitable set of features and/or arrangement.

In variants, a subset of the rotors (e.g., all rotors, rotors closest tothe ground in the forward configuration, rear rotors, etc.) can beretracted and/or captive within the nacelle in one or moreconfigurations, and actuated by the blade pitch mechanism and/or aretraction mechanism. In a first example, the rotors closest to theground retract during or after a fixed-wing style landing to protect therotors and/or protect humans from exposed blades.

The rotors are preferably unenclosed (e.g., without captive blade tips,without an inflow screen, without a fan duct, etc.), but in additionalor alternative variations can be enclosed (e.g., ducted as in a ductedfan, enclosed within a cowling about the perimeter of the disc area,etc.) and/or include a fixed screen in the inflow and/or outflow path.Rotor enclosures/ducting can be connected to the nacelle and/orotherwise mounted to the airframe.

The set of rotors can have any suitable arrangement on the aircraftand/or airframe. Rotors can be evenly or unevenly spatially distributedrelative to an axis of the aircraft, but can additionally or alternatelybe evenly or unevenly distributed relative to (e.g., about): theaircraft mass (e.g., about center of gravity), lift generation axes,aircraft geometry (e.g., an aircraft geometric center), airframegeometry (e.g., an airframe geometric center), and/or otherwisedistributed or arranged. The rotors are preferably symmetric about alateral plane of the aircraft and/or airframe, but can additionally oralternatively be asymmetric about the lateral plane. The rotors withinthe set can be: coplanar, offset (e.g., vertically, laterally,longitudinally, angled, etc.), or otherwise arranged relative to otherrotors in the set. In an example, rotors are evenly distributed on aleft and a right side relative to the longitudinal axis of the aircraftor airframe, but can be otherwise configured. Rotors, hubs, rotor discplanes (and/or swept area), and/or other suitable rotor references canbe: coplanar, offset from each other (e.g., in parallel planes), skewed(e.g., propeller axes can be skewed relative to one another), and/orotherwise configured in the forward, hover, transition, and/or otheroperation mode. In a first variant, distal (outboard) rotors arerecessed (arranged backwards, offset toward the aircraft rear in theforward configuration) from proximal (inboard) rotors, examples of whichis shown in FIGS. 14A-C and FIG. 17A-C. In a second variant, proximal(inboard) rotors are recessed (arranged backwards, offset toward theaircraft rear in the forward configuration) from distal (outboard)rotors, an example of which is shown in FIGS. 15A-C. In a third variant,inboard and outboard rotor disc planes are coplanar, an example of whichis shown in FIGS. 16A-C.

The rotors can have any suitable arrangement relative to the tilt axis112 and/or CoG in one or more modes of operation. One or more rotors(e.g., rotor hubs) can be: forward of the tilt axis, rearward of thetilt axis, above the tilt axis, below the tilt axis, arranged along thetilt axis, and/or otherwise suitably arranged relative to the tilt axisin the forward and/or hover configuration. One or more rotors (e.g.,rotor hubs) can be: forward of the CoG, rearward of the CoG, above theCoG, below the CoG, arranged along the lateral axis, and/or otherwisesuitably arranged relative to the CoG in the forward and/or hoverconfiguration. In a first example, a set of rotors (and/or correspondingrotor hubs) is arranged below the tilt axis and/or CoG (e.g., relativeto a vertical direction) in the forward configuration and above the tiltaxis and/or CoG in the hover configuration. In a second example, a setof rotors (and/or corresponding rotor hubs) is arranged forward of thetilt axis and/or CoG in the forward configuration and rearward of thetilt axis and/or CoG in the hover configuration.

The rotors can have any suitable arrangement relative to the wing.Rotors can be forward of the wing, rearward of the wing, above of thewing, below of the wing, arranged on an inboard portion of the wing,arranged on an outboard portion of the wing, within a boundaryprojection of the wing (e.g., wing functioning as a fan duct, etc.),and/or otherwise suitably arranged relative to the wing in the forwardand/or hover configurations. The rotors can have any suitablearrangement relative to the payload housing or fuselage. Rotors (and/orrotor hubs) can be above, below, forward, rearward, and/or within a sideview boundary projection of the payload housing or fuselage.

The set of rotors can define one or more rotor pairs, where rotorswithin each pair are arranged (e.g., entirely arranged, mostly arranged)on opposing sides of one or more axis of the aircraft (e.g., top/bottomrotor pair, left/right rotor pair, front/back rotor pair) in the hoverconfiguration and/or forward configuration. The aircraft can include anysuitable number of rotor pairs. The aircraft can include: 2, 3, 4, ormore than 4 pairs of rotors. In a specific example, the aircraftincludes three rotors per side: three left rotors and three rightrotors. However, rotor pairs can be otherwise defined, and rotors can beotherwise grouped. Examples of airframes including additional top/bottomand/or left/right rotor pairs are shown in FIGS. 18A-F and FIG. 19A-F.

In the forward configuration, the set of rotors preferably includes onerotor pair above and one rotor pair below the center of gravity (CoG) ofthe aircraft (and/or airframe or aircraft geometric center), but canadditionally or alternately include: more than one rotor pair aboveand/or below the CoG and/or airframe or aircraft geometric center (e.g.,2 above and 1 below, 2 above and 2 below), no rotor pairs above the CoGand/or airframe or aircraft geometric center, no rotor pairs below theCoG and/or airframe or aircraft geometric center, or be otherwisearranged. Preferably, one rotor pair is centrally aligned with the CoGin the forward configuration (e.g., for a maximally loaded aircraft, foran unloaded aircraft, within a specified range of CoGs) in a verticaldirection. Additionally or alternately, the aircraft can includemultiple rotor pairs aligned with the CoG, one rotor pair forward andone rotor pair of the lateral axis (e.g., the lateral axis intersectingthe CoG), one rotor of the rotor pair forward and one rotor of the rotorpair rearward of the lateral axis (e.g., the lateral axis intersectingthe CoG), and/or the aircraft can otherwise be suitably configured.

In a first specific example, the rotor hub and/or the lowest point ofthe rotor disc associated with a rotor mounted to the end of ananti-lateral support member extends below the base of the payloadhousing (e.g., cabin, fuselage, etc.) and/or extends below the landinggear (e.g., location of landing gear in forward configuration, lowestpoint of landing gear in landing configuration, etc.).

In the hover configuration, all rotor pairs preferably lie verticallyabove the CoG (loaded and/or unloaded) and/or aircraft geometric center,but additionally or alternately one or more rotors or rotor pairs can bearranged vertically below the CoG and/or aircraft geometric center. Inthe hover configuration, one rotor pair is preferably centrally alignedwith the CoG and/or airframe or aircraft geometric center in alongitudinal or vertical direction, but additionally or alternately:more than one rotor pair can be centrally aligned with the CoG and/orairframe or aircraft geometric center in a longitudinal or verticaldirection, at least one rotor pair can be forward and at least one rotorpair can be rearward of the CoG and/or airframe or aircraft geometriccenter, one rotor of a rotor pair can be forward and one rotor of therotor pair can be rearward of the lateral axis (e.g., the lateral axisintersecting the CoG or a geometric center) and/or rotors/rotor pairscan be otherwise arranged. However, the aircraft can be otherwisesuitably configured.

The set of rotors can include one or more unpaired rotors (e.g., for oddnumbers of rotors: 3, 5, 7, etc.). Unpaired rotors can be: centrallylocated relative to the tilt mechanism (e.g., lying in the in thesagittal plane defined by the longitudinal and vertical axes), locatedon the nose of the aircraft, located on a tail/trailing portion of theaircraft, and/or otherwise suitably located.

Preferably, in the forward configuration, all rotors and/or rotor pairsare arranged longitudinally forward of the tilt axis of the tiltmechanism, but alternately one or more rotors and/or rotor pairs liebehind the tilt axis. Preferably, in the hover configuration all rotorsand/or rotor pairs are arranged vertically above the tilt axis of thetilt mechanism, but alternately one or more rotors and/or rotor pairslie below the tilt axis. In a specific example, the tilt axis of thetilt mechanism lies above (along vertical axis) and behind (alonglongitudinal axis) the CoG of the aircraft.

In a first specific example: in horizontal (e.g., forward) flight, thecombined aerodynamic center of lift of all rotors (and/or rotor pairs)and/or airframe (e.g., including: wings, lift generated by the payloadhousing geometry, etc.) is substantially longitudinally aligned with theCoG of the aircraft (e.g., passing through the CoG; forming an angle of0 deg, <1 deg, <2 deg, <3 deg, <5 deg, <10 deg, and/or any othersuitable angle with the gravity axis or vertical axis; longitudinallyoffset from the CoG by an offset distance; etc.). In a second specificexample, a rear set of rotors connected to a tail of a payload housing(or payload housing pod) lie rearward of the tilt axis.

The rotors can define any appropriate thrust vector and/or lift vectorwith any suitable relationship to a drag vector and/or weight vector forthe aircraft in any mode of flight and/or configuration of the aircraft.Preferably, in the forward configuration (e.g., during horizontalflight), the forward thrust vector is substantially aligned with thedrag axis and the lift vector is substantially aligned with the weightvector (an example is shown in FIG. 6A), however the lift and/or forwardthrust vectors can alternately be offset, skewed and/or otherwiseoriented relative to the weight and/or drag vectors, respectively. Inthe hover configuration, the lift vector (or vertical thrust) vector issubstantially aligned with the weight vector (an example is shown inFIG. 6B), but can alternately be offset, skewed, and/or otherwiseoriented relative to the weight vector. In a specific example, the tiltaxis is offset from the center of lift and/or thrust and is arrangedoutside of the cabin area, while still providing a thrust vector that issubstantially aligned with the drag vector in forward flight and a liftvector that is substantially aligned with the CoG in hover. In a secondspecific example, a line extending from the CoG through the tilt axis inthe vertical/longitudinal plane defines a 45 degree angle relative tothe vertical and/or longitudinal axis. In a third specific example, thetilt axis can be arranged at a non-zero angle (e.g., 45 degrees) behindand above (and/or forward and below) the center of gravity, such thatthe average center of thrust of all rotors can still be aligned with thecenter of gravity in both hover and forward flight.

The rotors can rotate clockwise, counterclockwise, or a combinationthereof (e.g., wherein a subset of the rotors rotate clockwise and aremainder rotate counterclockwise). In operation, the rotors can rotatein the rotor's respective rotation direction at all times, switchrotation directions (e.g., based on the aircraft configuration, aircraftrotation or navigation, etc.), cease rotation, and/or otherwise operate.Half of the rotors preferably rotate in one direction, while the otherhalf counter rotate (e.g., in the other direction); however, the rotorscan be otherwise distributed between the two directions. At least onerotor per side (left/right) preferably rotates clockwise (and one rotorcounterclockwise), at least one rotor above a lateral support member(e.g., in the forward configuration) preferably rotates clockwise (andone rotor counterclockwise), and/or at least one rotor below a lateralsupport member rotates clockwise (and one rotor counterclockwise);however the aircraft can include two, more than two, zero, or anysuitable number of rotors in any of the aforementioned groups. Thedistribution of rotor rotation directions can be selected to enablecontinued flight operation and/or landing capability in the event of afailure of one or more propulsion assemblies. All rotors are preferablypowered in all modes of flight and/or flight configurations (e.g.,forward, transition, hover, etc.); however in variants a subset ofrotors can be unpowered during one or more modes of flight (e.g., duringforward flight, while changing heading, etc.)—which can conserve energyand/or improve heading control authority.

Example distributions of clockwise and counterclockwise rotors are shownin FIGS. 20A-D.

The support member of the aircraft 100, function to support the payloadhousing and transmit structural loads between the rotors. The set ofsupport members (or a subset therein) can, in variants, function togenerate lift in the forward configuration (e.g., during horizontalflight). In a first specific example, lateral support members generateless than a threshold proportion of aircraft lift in the forwardconfiguration, such as less than: 100%, 95%, 90%, 75%, 50%, 40%, 30%,20%, 10%, 5%, and/or any other proportion of aircraft (requisite) lift.In a second specific example, the lateral support members generatesubstantially no lift (e.g., less than 5%, less than 1%, etc.). In athird specific example, the lateral support members do not have anairfoil profile geometry over a portion or an entirety of the span. In afourth specific example, one or more support members of the set createsdrag or an aerodynamic force opposing lift or thrust.

The set of support members preferably includes multiple support members,but can additionally or alternatively include a single support member(e.g., a unitary piece). The set of support members can have anysuitable arrangement. Support members can have endpoints at: rotornacelles, tilt mechanisms, the payload housing exterior, other supportmembers, and/or other components on the aircraft. The support memberspreferably connect endpoints rigidly (e.g., mechanically bonded,integrated, fastened, etc.), but alternately one or more endpoints canbe rotatably connected (e.g., pivot relative to the tilt mechanism,pivot relative to the payload housing, pivot relative to a remainder ofthe airframe, etc.) and/or otherwise connected. Support members canconnect any number of endpoint nodes on the aircraft (e.g., 2 endpointnodes). In one example, the support member set preferably includes atleast 2 anti-lateral support members and 1 lateral support member.However, the set of support members, can include any number supportmembers in any suitable configuration.

In a first variant—the dual quadrilateral variant—the support memberscooperatively form a closed quadrilateral geometry on each side of theaircraft (e.g., mirrored about the sagittal plane), with the threecorner nodes of the quadrilateral lying at rotor nacelles and the fourthconnecting them to the tilt mechanism or an extension off of the tiltmechanism (e.g., by an additional member). A specific example of thedual quadrilateral variant is shown in FIG. 4D. A second specificexample of the dual quadrilateral variant is shown in FIG. 4H.

In a second variant—the bowtie variant—the support members form an outerpolygon with corners at each of the rotor nacelles (examples of polygongeometries connecting different numbers of rotors shown in FIGS. 5A-5F),and a lateral support connects the distal (in the lateral direction)rotor nacelles to the tilt mechanism. In a first example, the lateralsupport span as a single continuous member. In a second example, thelateral support is divided across the payload housing and connected tothe tilt mechanism at either side. The tie variant includes two opposinglateral support members on opposite sides of the payload housing whichform the upper and lower members of the outer polygon in the forwardconfiguration, and the forward and rearward members of the outer polygonin the hover configuration. The distance between the two opposinglateral support members can exceed (e.g., by a clearance distance suchas 50 cm), both the length and height of the payload housing to avoidimpacting the payload housing during the transition between forward andhover, but can be otherwise dimensioned. A specific example of the tievariant is shown in FIG. 4B.

In a third variant—the spaceframe variant—a different support memberextends between each nacelle and the tilt mechanism (e.g., for N rotors,N support members connect the N rotors to the tilt mechanism). Anadditional lateral support connects the two uppermost (in the forwardconfiguration) nacelles, and an additional set of support members formsa closed polygon with each of the rotor nacelles on the same side of theaircraft as endpoints (e.g., for six rotors: triangle on the left sideand right side). A specific example of the spaceframe variant is shownin FIG. 4A.

In a fourth variant—the cross variant—the cross variant includes a setof lateral support members spanning between the left outermost and rightoutermost nacelles, and anti-lateral support members connecting aremainder of the rotors to the set of lateral support members. The setof lateral support members can be a single spanning lateral supportmember (an example is shown in FIG. 4E) or can include a left member anda right member connecting on either side of the tilt mechanism (anexample is shown in FIG. 4G). Alternately, the endpoint on each side ofthe lateral support member can connect to another anti-lateral supportmember (e.g., for 8 rotor example shown in FIG. 5F).

In a fifth variant—the cross-beam variant—the nacelles are joined as inthe cross variant, with the upper nacelles (in the forwardconfiguration) connected by an additional support. This support canconnect to the tilt mechanism, the set of lateral support members (fromthe cross variant), or to a beam extending from the tilt mechanism. Anexample of the cross-beam variant is shown in FIG. 4F.

In a sixth variant, the aircraft can include 2, 3, 4, 5, 6, 8, 10, 12,14, and/or any other suitable number of rotors arranged as described inone of the aforementioned variants (e.g., with additional left/right,top/bottom, and/or front/back rotor pairs) and/or otherwise suitablyarranged.

The set of support members can have any suitable angle(s) relative tothe payload housing, tilt mechanism, other support members, and/or otherreference point. The support members can extend perpendicular relativeto the sagittal plane (straight out), or define an anhedral angle (anexample is shown in FIG. 10B and FIG. 19A-F), dihedral angle (an exampleis shown in FIGS. 10A and 18A-F), forward sweep angle, rearward sweepangle. The angles can be: 0 deg, 5 deg, 10 deg, 15 deg, 20 deg, 30 deg,45 deg, 60 deg, 70 deg, 75 deg, 80 deg, 85 deg, 90 deg, 0-15 deg, 15-30deg, 30-60 deg, 60-75 deg, 75-90 deg, and/or any other suitable angle.The shape of the support members and/or airframe structure can be: aninverted gull shape, a gull shape, dihedral, anhedral, and/or any othersuitable shape or geometry.

The set of support members can be assembled together in any suitable wayto form the airframe (or a portion of the airframe, such as the supportstructure), and examples of this are shown in FIGS. 4A-4H. Preferably,the support members are arranged such that they have non-intersectingtorsion boxes 148, such that the stiffness is not diminished by anirregularity in the internal structure/geometry, however the supportmembers can alternately intersect (an example is shown in FIG. 8C), havenesting geometry exterior geometry, and/or be otherwise arranged.Preferably, anti-lateral support members are arranged forward of lateralsupport members (an example is shown in FIG. 8A), but can additionallyor alternately be rearward of lateral support members (an example isshown in FIG. 8B). Connections between two support members can have anyappropriate geometry, they can be: filleted, arcuate with constantradius, arcuate with variable radius, chamfered, aerodynamicallyoptimized, and/or otherwise connected. The connection can be integratedwith the support member (e.g., same component), a different component ofthe same material (e.g., composite), or different material (e.g.,aluminum gusset fastened or bonded to the torsion boxes). In variantsutilizing composite structures, the components can be manufactured inthe same composite layup, manufacturing can include different compositelayups for external support structures, or the components can beotherwise manufactured/bonded. In variants, more than two supportmembers can connect and/or intersect at a single endpoint, node, orconnection.

In a first variant, torsion boxes 148 can occupy a portion of the crosssection of the support member cross-sectional profile, such thatintersecting support members can have non-intersecting torsion boxes. Anexample is illustrated in FIG. 22A.

The support members can have any exterior geometry. Support memberspreferably have a symmetric airfoil cross section, but can alternatelyhave an asymmetric airfoil cross section; an aerodynamically efficientshape, such as a rounded/arcuate leading edge and a tapering thicknesstowards a trailing edge (e.g., pointed, flat, etc.); a stiffness-drivengeometry (e.g., circular cross section, hexagonal cross section, etc.);a non-aerodynamic geometry; rectangular, ovoid, circular, triangular,and/or any other suitable cross section geometry. Support members canbe: straight, constant cross section, arcuate (e.g., curved, bowed),tapered (e.g., reduction in cross sectional area along the span in adirection of the taper), un-tapered (e.g., constant cross sectionalprofile), gull-wing (or inverted gull wing) shaped, angled, and/or haveany other suitable spanwise shape. Lateral support members can bestraight, bent, and/or curved along a longitudinal axis, and/or defineany other suitable geometry.

In a specific variant, support members can be wings 145, and caninclude: a left wing, a right wing, a top wing, front wing, rear wing, abottom wing, a wing spanning a full width of the airframe.

The set of support members can have any suitable internal construction.Preferably, support members include a torsion box which is integrated orconnected internally inside of the support member to provide torsionalstiffness/rigidity. The torsion box has a tubular structure can have thesame cross-sectional profile as the support member (e.g., ribs and sparssupporting the exterior), or have a different cross-sectional profilesuch as: round, square, hexagonal, rectangular, and/or any othersuitable geometry. Preferably, the torsion box has an internal grid ofbeams/supports extending in a radial, lateral, longitudinal, vertical,skew, or other orientation, but can also have a honeycomb, 3D honeycomb,prismatic, wave pattern, rib/spar, and/or other internal structure.However, a support member can be: solid (e.g., solid beams), hollowshells, include internal trusses, or otherwise constructed. The supportmembers (and internal structures) are preferably made of a compositematerial (e.g., carbon fiber, fiberglass, etc.), but can additionally oralternately include metal or metal alloys (e.g., steel or aluminum),plastic, any combination thereof, and/or any other suitable material.

The set of support members can include any suitable number of lateralsupport members 142. The set of support members can include: 1, 2, 3, 4,5, 6, or more than 6 lateral support members. Preferably, “lateral”support members as referenced herein refer to support members whichdefine an angle greater than 45 degrees with respect to the sagittalplane of the aircraft or airframe (i.e. perpendicular to the sagittalplane or closer to perpendicular than parallel). However, a lateralmember can additionally or alternately be: support members that definean angle with the sagittal plane which is greater than: 45 deg, 50 deg,60 deg, 75 deg, 80 deg, 85 deg, support members exactly perpendicular tothe sagittal plane, support member exactly parallel to the frontalplane, support members, and/or otherwise suitably defined. The set ofsupport members can include any suitable number of lateral supportmembers per side, such as one per side, two per side, three per side,and/or any other suitable number. The set of support members can includea rigid connection between the wings on opposing sides of the aircraft,or the left and right sides can be split (for example, the left andright wings can be independently actuated by the tilt mechanism; anexample is shown in FIGS. 24A-B).

The lateral support members can individually or collectively define anysuitable wing area for the aircraft. The wing area can refer to: thevertical projection of the wing, the chord length integrated in thespanwise direction, the area of the upper surface of the wing, and/orotherwise suitably defined. However, the wing area can otherwise besuitably defined.

The set of support members can include any suitable number ofanti-lateral support members 144 (e.g., vertical support members). Theset of support members can include: 1, 2, 3, 4, 5, 6, or more than 6anti-lateral support members. Preferably, anti-lateral support membersas referenced herein refer to support members which define, in theforward configuration, an angle less than 45 degrees with respect to thesagittal plane (i.e. parallel to the sagittal plane or closer toparallel than perpendicular); however the threshold to define ananti-lateral support member can additionally or alternately be supportmembers which define an angle with the sagittal plane which is lessthan: 5 deg, 10 deg, 15 deg, 20 deg, 30 deg, 40 deg, support membersexactly parallel to the sagittal plane, support members perpendicular toa lateral support member, any support member which is not a lateralsupport member, and/or otherwise suitably defined. The sagittal plane(e.g., longitudinal plane) can extend along: the payload housing'slongitudinal and vertical axis, the wing or lateral support member'sspan and chord, or be otherwise defined. An anti-lateral support member144 is preferably mounted to a lateral support member, but can beotherwise mounted. Each anti-lateral support member preferably extendsbeyond a first and second side (e.g., broad face, broad surface) of therespective mounting component (e.g., above and below the lateral supportmember in the forward configuration), but can additionally oralternatively: extend beyond a single side of the mounting component,include a first and second end arranged on opposite sides of themounting component's broad surface or transverse plane (e.g.,encompassing a span and chord of the lateral support member), or beotherwise arranged.

In a specific variant, the aircraft includes a lateral support member(e.g., wing) and an anti-lateral support member mounted to and fullysupported by the lateral support member. In the specific variant, one ormore rotors can be mounted to the lateral support member outboard of thelateral support member, mounted to the lateral support member inboard ofthe anti-lateral support member, mounted to the anti-lateral supportmember (above and/or below the connection between the anti-lateralsupport member and the lateral support member), mounted to the end ofthe anti-lateral support member, mounted to the end of the lateralsupport member, and/or otherwise arranged.

In a first example of the specific variant, the airframe includes: aleft wing; a right wing; a left anti-lateral support member coupled tothe left wing; and a right anti-lateral support member coupled to theright wing. A first end of the left anti-lateral support member isarranged above the left wing and a second end of the left anti-lateralsupport member is arranged below the left wing. A first end of the rightanti-lateral support member is arranged above the right wing and asecond end of the right anti-lateral support member below the rightwing. In the first example of the specific variant, the airframe furtherincludes: a left outboard propulsion assembly mounted to the left wingoutboard of the left anti-lateral support member; a right outboardpropulsion assembly mounted to the right wing outboard of the rightanti-lateral support member; a first and second inboard propulsionassembly mounted to the first and second ends of the left anti-lateralsupport member, respectively; and a third and fourth inboard propulsionassembly mounted to the first and second ends of the right anti-lateralsupport member, respectively; wherein each of the plurality ofpropulsion assemblies comprises: an electric motor; and a propellerrotatably coupled to the electric motor about an axis of rotation. Theleft wing can include a torsion box extending from the tilt mechanism tothe left outboard propulsion assembly, and the left anti-lateral supportmember can include an anti-lateral torsion box extending from the firstpropulsion assembly to the second propulsion assembly. Preferably, theanti-lateral torsion box does not intersect the torsion box in the wing;however the anti-lateral torsion box can alternately intersect thewing's torsion box, be formed as a single component, and/or be otherwiseimplemented. The torsion box in the left wing can terminate at the tiltmechanism (an example is illustrated in FIG. 22B), terminate at anacelle node of a propulsion assembly, extend from the left outboardpropulsion assembly to the right outboard propulsion assembly (anexample is shown in FIG. 22C), or be otherwise configured.

The set of support members can optionally include any suitable number ofground support members. Ground support members function to support theaircraft on the ground (and/or in a taxi configuration). Ground supportmembers can be integrated into the airframe, connected to the tiltingmechanism, connected and/or integrated into the rotor nacelles, and/orotherwise implemented. Preferably, ground support members extend to theground (e.g., below the payload housing) in the hover configuration ofthe tilt mechanism, but can alternately always extend below the payloadhousing (e.g., connected to the same part of the tilting mechanism asthe payload housing), and/or otherwise support the aircraft on theground. Ground support members can optionally be dampened (e.g., withrubber, compressible springs, etc.) and/or include rollers (e.g., forground taxiing), but can alternately include skids (e.g., for waterlandings) and/or be otherwise implemented. Preferably, ground supportmembers contact the ground at three or more points (e.g., 4 points, 6points, for N number of rotor nacelles contact at N points, etc.), butcan be otherwise configured.

The aircraft 100 can include a tilt mechanism which functions totransform the rotors between the forward configuration and the hoverconfiguration. The tilt mechanism can optionally adjust the angle ofattack, change the dihedral angle, or otherwise actuate the supportmember and/or rotor position. The tilt mechanism is preferablyelectromechanical, but can alternately operate by pneumatic, hydraulic,and/or other actuation. The tilt mechanism actuation can be achieved bya rotation, linear actuation, combination of rotary and linearactuation, or otherwise actuated. The tilt mechanism is preferablyintegrated into a support member (e.g., middle section of a lateralsupport member), but can alternately mount to a support member, mount tomultiple support members, mount to the airframe, mount to the payloadhousing, be integrated with the payload housing, and/or be otherwiseimplemented. When the tilt mechanism mounts multiple support members,the tilt mechanism can actuate the multiple support members:independently, together, and/or with any other suitable relationship.Support members connected to the tilt mechanism can be: cantilevered,over-hanging, double over-hanging, trussed (by additional supportmembers/connections), and/or otherwise connected. The payload housing(and/or cargo pod) is preferably suspended from the tilt mechanism, butcan additionally or alternately be integrated into, bonded, fastened,and/or otherwise connected to the tilt mechanism. The tilt mechanism canbe incorporated into the airframe, with the pod attaching to a residualpart of the airframe which separate from the wing, wherein the tiltmechanism actuation is fully contained within the airframe.Alternatively, the tilt mechanism can be separate from the airframe. Ina first variant, the tilt mechanism is connected to the payload housingand a middle section of a lateral support member. In a second variant, aleft side of the tilt mechanism connects to the end of a left lateralsupport member, and a right side of the tilt mechanism connect to theend of a right lateral support member. In a first example of the secondvariant, the tilt mechanism connects to the top of the payload housing.In a second example of the second variant, the tilt mechanism isintegrated into the exterior of the payload housing (and/or fuselage).

In variants, the tilt mechanism can transform the wing (or lateralsupport members) by a transformation angle 194, which can be >95 deg, 95deg, 92 deg, 90 deg, 89 deg, 87 deg, 85 deg, 80 deg, 75 deg, <75 deg,any range bounded by the aforementioned angles, and/or any othersuitable angle.

The tilt mechanism can have any suitable arrangement on the aircraftrelative to the CoG, principal axes (e.g., lateral, longitudinal,vertical), and/or any other suitable reference. The tilt mechanism ispreferably centered along the lateral axis of the aircraft, symmetricabout the sagittal plane of the aircraft, and/or otherwise locatedlaterally on the aircraft. The tilt mechanism preferably defines a tiltaxis about which the tilt mechanism pivots/rotates. The tilt axis ispreferably above and/or behind the: aircraft CoG (an example is shown inFIG. 23A-B), CoG of the payload, passenger region (an example is shownin FIGS. 11A-B), payload housing, and/or other reference on theaircraft. Additionally or alternately, the tilt axis can be centeredabout, forward of, and/or below the lateral (pitch) axis and/or the CoG,or otherwise suitably located. In a first variant, the tilt axis islocated on an upper portion of the payload housing, above the payloadhousing, and/or otherwise does not infringe on the payload housingspace. In a second variant, the tilt axis extends through the thicknessof the payload housing wall, but does not extend into a portion of thepayload housing where passengers and/or cargo reside. In a thirdvariant, the tilt axis lies above the payload housing.

In a first specific example, a left component of the tilt mechanism anda right component of the tilt mechanism pivot a left set of the supportmembers and a right set of the support members relative to the payloadhousing about the tilt axis (an example is shown in FIG. 10B),respectively.

In a second specific example, a left component of the tilt mechanism anda right component of the tilt mechanism pivot a left set of the supportmembers about a first tilt axis and a right set of the support membersabout a second tilt axis, respectively, wherein the first tilt axis andthe second tilt axis lie in the same plane.

The tilt mechanism can optionally operate in conjunction with a lockingmechanism which functions to prevent the payload housing, airframe,support members, rotors, and/or other components from rotating in anuncontrolled or unintended manner if the tilt mechanism fails.Preferably, the locking mechanism defaults to a locked position (e.g.,in power failure scenario) which does not require continuous power toretain the angular position of the tilt mechanism. The locking mechanismcan include: a non-backdrivable mechanism (e.g., ratcheting, worm-ger,etc.), hydraulic locking, pneumatic locking, an external brakingmechanism (e.g., such as a disk brake), and/or other locking mechanism.The locking mechanism can engage: in a power failure scenario, if therelative angular position of the left side and the right side of theairframe about the tilt axis exceeds a predetermined threshold (e.g., 1deg, 3 deg, 5 degrees, etc.), in response to a user input, and/or on anyother event driven basis.

The tilt mechanism can additionally function to transform the lateralsupport members (e.g., left wing, right wing, wings, etc.) about thepitch axis (e.g., lateral axis). The tilt mechanism can rotate thelateral support members by >95 deg, 95 deg, 92 deg, 90 deg, 89 deg, 87deg, 85 deg, 80 deg, 75 deg, <75 deg, any range bounded by theaforementioned angles, and/or any other suitable angle.

The tilt mechanism can additionally or alternately operate inconjunction with one or more rotor tilt mechanisms, which can pivot aset of rear rotors, outboard rotors, and/or other rotors independentlyof the remainder of the airframe (e.g., about an axis different from thetilt axis). In a specific example, the rotor tilt mechanism is themechanism described in U.S. application Ser. No. 16/409,653, filed 10May 2019, which is incorporated in its entirety by this reference.However, any other suitable tilt mechanism can be used.

In a first variant, the left and right wings are fixed relative to eachother. In a first example, a torsion box extends through the tiltmechanism and couples the left wing to the right wing. In a secondexample, each wing includes a torsion box rigidly connected at the tiltmechanism. In a third example, the tilt mechanism can include a leftactuator connected to the left wing, a right actuator connected to theright wing, and an interlock preventing relative motion of the left andright actuators (e.g., beyond a threshold, etc.). The interlock can bedefault locked or default unlocked, can be passive or active, can bemechanical or electromechanical, or can be any other suitable type ofinterlock.

In a second variant, the left and right wings can be independentlytransformed relative to the pitch axis, which can enable additional rollcontrol authority (e.g., no bank turns, tighter turn radius, etc.). Thiscan present an additional failure mode—loss of tilt control function ofone or both wings—which can be mitigated by additional redundancies inthe system, such as for aircrafts with human passengers or triplyredundant aircrafts. In less safety critical aircrafts (e.g., unmannedaircraft, autonomous aircraft, delivery aircraft, etc.), loss of tiltcontrol can be mitigated by control augmentation, emergency landing,and/or otherwise mitigated or unmitigated.

In a third specific variant, the tilt mechanism of the aircraft canchange the aerodynamic force generated by the wing(s) and the net thrust(and/or lift) from the rotors by independently tilting the left and/orright wings of the aircraft (and all of the propulsion assembliesmounted thereto). Independent actuation of the left and right wings cancreate a net yaw moment while balancing roll moments, thereby enablingheading changes without banking the aircraft.

The aircraft can include a payload housing coupling mechanism, whichfunctions to connect the payload housing to the tilt mechanism. Thepayload housing coupling can include mechanical which connect thepayload housing to the tilt mechanism, support member structure, and/orother aircraft structural elements. The payload housing coupling caninclude electrical connections which connect: sensors, pilot controls,HVAC, lighting, and/or other equipment requiring electrical connectionsin the payload housing. The electrical connections can include wiremanagement (e.g., slip ring connector or similar) to avoid stressing,fatiguing, and/or damaging wires during the transition between forwardand hover. In variants including a modular payload housing pod, thepayload housing coupling mechanism can selectively connect anddisconnect the payload housing from the airframe, a remainder of theaircraft structure, and/or various electrical endpoints (e.g.,batteries, motors, etc.).

The aircraft 100 can include a payload housing which functions toprotect and carry the aircraft payload. The payload can include: 1 ormore human passengers (e.g., 2) and/or articles of luggage, packages,cargo, food deliveries, and/or other equipment related to aircraftoperation. Preferably, the payload include 3 or more human passengers(e.g., 3, 4, 5, 6, more than 6), and can optionally include a pilot(along with pilot input control mechanisms). Additionally oralternately, the aircraft can be remotely piloted and/or operateautonomously.

In a first variant, the aircraft is a delivery drone, and the payloadhousing carries packages such as food or user goods directly to a useror intermediary (e.g., service location, vehicle, distribution center,delivery person, etc.).

In a second variant, the aircraft is a camera drone. The payload housingcarries a camera used for scanning a region and/or performing imagingservices. The aircraft can optionally be equipped with onboard memoryfor storing the images and/or can be equipped to stream imaging data toa user or remote system.

In a third variant, the aircraft is employed for aerial dispersion. Thepayload housing can be equipped to carry a chemical agent (e.g.,pesticide), animals (e.g., living fish), organic particulates (e.g.,seeds, soil), water, and/or other payloads. The payload housing canoptionally include a dispersion system for ejecting the payload via apressurized or aerosol spray, release hatch, or other means ofdispersal. The cabin, cargo hold, and/or payload housing can: includeinsulation 122 (such as in the example in FIG. 21) or be uninsulated,include temperature conditioning (e.g., via heating and/or coolingsystems) or be unconditioned (e.g., no onboard heating and/or coolingsystems), be air-tight or not air-tight, include windows or not includewindows (e.g., or otherwise not optically connected to the aircraftexterior), include an inceptor (or other pilot input mechanism) or notinclude an inceptor (or other pilot input mechanism), include cargobays, include tiedowns, house a battery pack or other power source,and/or include any other suitable characteristics/features. The payloadhousing can optionally include a cover 121 which can function toselectively allow access to the interior, such as during payload loadingand/or unloading. The cover can be a clamshell, gull-wing, side-open(e.g., like a car door), sliding, snap in, hinged, bottom panel, and/orother suitable cover. The cover can additionally or alternately functionto form a portion or entirety of the external profile of the payloadhousing, and to enclose and/or protect the payload (e.g., passengers,delivery goods, etc.).

In a specific variant, the payload housing includes a modular pod whichfunctions to detach from the remainder of the aircraft to allowmodularity and reconfigurability. The pod can connect and/or disconnect:automatically, partially automatically (e.g., pilot initiated sequence),manually, and/or in any other suitable manner. In a specific example,the pod is connected with the assistance of an alignment mechanism (anexample method is shown in FIG. 9), which can operate based on: computervision; physical alignment features such as tapering grooves/channels, areference frame or key, and/or other self-locating geometry;variance-tolerant connections (e.g., can tolerate: <1 mm deviation, <5mm deviation, <1 cm deviation, <3 cm deviations, <5 cm deviations,etc.); and/or any other suitable mechanisms. The payload housingcoupling mechanism which operates in conjunction with a modular pod canutilize any suitable mechanical fasteners and/or mechanical couplingtechnique (e.g., to achieve clamping, retaining, latching, etc.) suchas: magnetic fasteners, hydraulic actuators, pneumatic actuators,electromechanical actuators, spring assisted actuators, and/or otherwisefastening or securing the modular pod.

The aircraft 100 can optionally include an impact attenuator whichfunctions to mitigate effects of an impact on the aircraft and/orpayload. The impact attenuator can be located at the rear of the pod (anexample is shown in FIG. 11A), below the passenger region, in front ofthe passenger region, at the front of the payload housing, at the bottomof the payload housing, side of the payload housing, and/or in any otherappropriate location. The impact attenuator can be mounted internally,externally, integrated into the construction of the airframe or payloadhousing, and/or otherwise mounted. In a first variant, the impactattenuator is passive: one or more portions of the payload housing arecrushable, collapsible, and/or deformable to mitigate theforces/accelerations experienced by the payload (e.g., passengers) in animpact scenario. The crushable portions of the aircraft can be the sameor different material from a remainder of the payload housing. They canbe constructed of: foam, aluminum (e.g., in a honeycomb structure),spring steel, and/or other material. In a second variant, the impactattenuator can include a self-inflating gas cushion (e.g., airbag) orother active safety system which deploys on impact.

The aircraft 100 can optionally include a set of power sources whichfunction to supply power to the set of rotors. Preferably, the powersource includes one or more batteries (e.g., arranged into batterypacks), but can additionally or alternately include fuel cells, liquidfuel (e.g., gasoline, diesel, jet fuel, etc.), and/or any other suitablepower sources. The power sources can be housed: inside nacelles, in theairframe, outside of the payload housing, inside of the payload housing,in an empennage, inside of support members, mounted to support members,in a dedicated enclosure mounted to the airframe, and/or in any otherappropriate location.

In variants, the power source can be hot swappable, removable,replaceable, and/or interchangeable. In a specific example—where thepower source includes one or more battery packs—battery packs can beconfigured to be charged onboard the aircraft, configured to be chargedoffboard the aircraft (e.g., removed for charging), configured to bereplaced with a fully charged battery pack, and/or otherwise configured.Such variants can increase aircraft uptime, reduce charging time, reducerequired part count on the aircraft, and/or otherwise improveoperational efficiency.

The airframe preferably includes the nacelles, support members (and theconnections between the support members), and/or any other structuralcomponents. In a first specific example, the airframe can include astructural payload housing with the tilt mechanism integrated into thepayload housing. In a second specific example, the airframe includes anempennage (e.g., with a set of control surfaces, with a set of rearrotors, etc.). An example of an aircraft including an empennage is shownin FIGS. 12A-D.

The aircraft 100 can include various flight control elements tofacilitate flight control and operation, which can include controlsurfaces and/or control actuators. For example, the aircraft 100 caninclude landing gear (e.g., retractable landing gear, powered/unpoweredwheels, fixed landing gear, nacelle struts, etc.), flight controlsurfaces (e.g., flaps, elevators, ailerons, rudders, ruddervators,spoilers, slats, air brakes, etc.), flight instruments (e.g., altimeter,airspeed indicator and measurement device, vertical speed indicator andmeasurement device, compass, attitude indicator and measurement device,heading indicator and measurement device, turn indicator and measurementdevice, flight director systems, navigational systems, and any othersuitable instruments), and any other suitable components. The variouscomponents can be coupled to the aircraft 100 in any suitable manner;for example, the flight control surfaces can be coupled to and/ordefined by portions of the airframe and/or the tail; the flightinstruments can be arranged within a payload housing (e.g., cockpit) ofthe aircraft 100 and/or at a remote operation location (e.g., ateleoperation facility, a remote piloting location, etc.); or otherwisearranged.

In variants, the payload housing of the aircraft (e.g., a payloadhousing pod, payload housing built into the airframe, etc.) can includean empennage, including any suitable set of flight control surfaces.

The aircraft can optionally include an empennage which functions tostabilize the aircraft by balancing aerodynamic moments in the forwardconfiguration. The empennage can include a set of stabilizers 123, whichcan include lateral stabilizers (e.g., a rear wing) and/or verticalstabilizers (e.g., tail fin). The lateral stabilizer preferably does notinclude an elevator (or ruddervator) or other actuator, but canalternately include an elevator or other control surface. The verticalstabilizer or does not include a rudder or other actuator, but canalternately include a rudder or other control surface. Eliminatingcontrol surfaces from the vertical and/or lateral stabilizers can reduceweight, reduce the number failure modes, reduce aircraft complexity(e.g., total part count), improve manufacturability, and/or confer anyother suitable benefits. Stabilizers can be integrated into the payloadhousing/cabin, mounted to the payload housing, or otherwise suitablyconnected to the aircraft. In variants, the empennage can be selectivelyattachable/detachable (e.g., selectively attached in high windconditions, etc.), can be formed with and/or manufactured by the sameprocess as the payload housing, can be manufactured separately (of sameor different material as the payload housing) and connected to thepayload housing, and/or otherwise mounted.

Embodiments of the system and/or method can include every combinationand permutation of the various system components and the various methodprocesses, wherein one or more instances of the method and/or processesdescribed herein can be performed asynchronously (e.g., sequentially),concurrently (e.g., in parallel), or in any other suitable order byand/or using one or more instances of the systems, elements, and/orentities described herein.

As a person skilled in the art will recognize from the previous detaileddescription and from the figures and claims, modifications and changescan be made to the preferred embodiments of the invention withoutdeparting from the scope of this invention defined in the followingclaims.

We claim:
 1. An electric aircraft system comprising: a payload housing;an airframe coupled to the payload housing, the airframe comprising: aleft wing defining a first broad surface; a right wing defining a secondbroad surface; a left support member mounted to the left wing, the leftsupport member comprising a first end and a second end arranged onopposite sides of the first broad surface; and a right support membermounted to the right wing, the right support member comprising a firstend and a second end arranged on opposite sides of the second broadsurface; a tilt mechanism rotatably coupling the left wing and the rightwing to the payload housing, the tilt mechanism configured to transformthe electric aircraft system between a forward configuration and a hoverconfiguration by rotating the left wing and right wing between a firstand a second position; and a plurality of propulsion assembliescomprising: a left outboard propulsion assembly mounted to the left wingoutboard of the left support member; a right outboard propulsionassembly mounted to the right wing outboard of the right support member;a first and second inboard propulsion assembly mounted to the first andsecond ends of the left support member, respectively; and a third andfourth inboard propulsion assembly mounted to the first and second endsof the right support member, respectively, wherein each of the pluralityof propulsion assemblies comprises: an electric motor; and a propellerrotatably coupled to the electric motor about an axis of rotation at afixed angle of attack relative to a wing chord line, the fixed angle ofattack defined between the axis of rotation and the wing chord line,wherein the fixed angle of attack is non-zero.
 2. The electric aircraftsystem of claim 1, wherein the fixed angle of attack is between 3degrees and 9 degrees.
 3. The electric aircraft system of claim 1,wherein the left wing comprises a torsional stiffening member extendingfrom the tilt mechanism to the left outboard propulsion assembly,wherein the left support member comprises an anti-lateral torsionalstiffening member extending from the first propulsion assembly to thesecond propulsion assembly, wherein the anti-lateraltorsional-stiffening member does not intersect the torsional-stiffeningmember.
 4. The electric aircraft system of claim 1, wherein thepropeller of each propulsion assembly defines a disc area and a discplane containing the disc area, wherein the propeller of each propulsionassembly comprises a hub at the center of the disc area, wherein the hubof the second and fourth inboard propulsion assemblies is below a baseof the payload housing in the forward configuration.
 5. The electricaircraft system of claim 4, wherein the electric aircraft system definesa center of mass, a lateral axis extending through the center of mass, avertical axis extending through the center of mass, and alateral-vertical plane containing the lateral axis and the verticalaxis, wherein the hub of the first inboard propulsion assembly and thehub of the left outboard propulsion assembly are arranged on opposingsides of the lateral-vertical plane in the forward configuration.
 6. Theelectric aircraft system of claim 5, wherein the tilt mechanism isconfigured to rotate the left and right wing about a tilt axis, whereinthe tilt axis is rearward of the center of mass.
 7. The electricaircraft system of claim 5, wherein the left wing and the right wing areanhedral.
 8. The electric aircraft system of claim 1, wherein the tiltmechanism is configured to rotate the left and right wings less than 90degrees between the forward and hover configurations.
 9. The electricaircraft system of claim 1, wherein the electric aircraft system doesnot comprise an elevator, aileron, or rudder.
 10. The electric aircraftsystem of claim 1, wherein the electric aircraft system defines a weightvector and is configured to generate a net lift vector opposing theweight vector during flight, wherein the propulsion assemblies areconfigured to generate at least 25 percent of the net lift vector duringforward flight.
 11. The electric aircraft system of claim 1, wherein theelectric aircraft system is configured to change a heading of theelectric aircraft system during forward flight without banking.
 12. Theelectric aircraft system of claim 1, wherein the propulsion assembliesare configured to simultaneously regenerate electrical energy andgenerate a net moment about the aircraft center of mass.
 13. Theelectric aircraft system of claim 1, wherein the left and right wingscooperatively define a total projected wing area, wherein the propellersof the propulsion assemblies cooperatively define a total projectedblade area, wherein the total projected blade area is between 50% and200% of the total projected wing area.
 14. The electric aircraft systemof claim 1, wherein each propeller of the plurality of propulsionassemblies comprises five blades.
 15. An electric aircraft systemcomprising: a payload housing; a left wing defining a first broadsurface; a right wing defining a second broad surface; a left supportmember mounted to the left wing, the left support member comprising afirst end and a second end arranged on opposite sides of the first broadsurface; and a right support member mounted to the right wing, the rightsupport member comprising a first end and a second end arranged onopposite sides of the second broad surface; a tilt mechanism connectingthe left wing and the right wing to the payload housing, the tiltmechanism configured to independently rotate the left wing and the rightwing relative to the payload housing; and a plurality of propulsionassemblies comprising: a left outboard propulsion assembly mounted tothe left wing outboard of the left support member; a right outboardpropulsion assembly mounted to the right wing outboard of the rightsupport member; a first and second inboard propulsion assembly mountedto the first and second ends of the left support member, respectively;and a third and fourth inboard propulsion assembly mounted to the firstand second ends of the right support member, respectively.
 16. Theelectric aircraft system of claim 15, wherein the left wing comprises atorsional stiffening member extending from the tilt mechanism to theleft outboard propulsion assembly, wherein the left support membercomprises an anti-lateral torsional stiffening member extending from thefirst propulsion assembly to the second propulsion assembly, wherein theanti-lateral torsional-stiffening member does not intersect thetorsional-stiffening member.
 17. The electric aircraft system of claim15, wherein the electric aircraft system is configured to change aheading of the electric aircraft system during forward flight withoutbanking.
 18. The electric aircraft system of claim 17, wherein theelectric aircraft system does not comprise an elevator, aileron, orrudder.